The present invention relates in general to techniques for manufacturing composite-material structures.

Owing to the increasingly stringent certification requirements and the need to be competitive in the aeronautical market in terms of costs, weight and performance, innovative aeronautical structures based on the use of composite materials are becoming necessary. Liquid resin infu- sion performed outside of an autoclave provides the basis for overcoming the shortcomings preventing the use of latest generation low-cost construction technology since it allows a greater integration of large structural components compared to the technology which uses pre- impregnated materials. The existing processes and materials, in fact, since they are precisely designed and developed for large-size military or civil transportation applications, result in the need for a series of standards, measures, steps and checks which make them incompatible with low-cost applications and with small production volumes. Demonstrating the feasibility of realizing highly integrated structures results in technology which is more convenient from the point of view of cost containment. Furthermore the use of an oven instead of an autoclave is able to ensure a two-thirds reduction of capital costs. A further reduction in costs is possible in that the cost of the dry fibre and liquid resin may be up to 30 percent lower compared to that of the same materials converted into prepregs.

One object of the present invention is therefore to provide a method for manufacturing integrated composite-material structures based liquid resin infusion techniques.

According to the present invention the aforementioned object is achieved by means of a method for manufacturing a composite-material panel structure, comprising the following steps:

a) forming, on a first unit, a dry-fibre laminate of the panel, called skin, comprising a plu- rality of layers;

b) forming or arranging on a plurality of support blocks a plurality of transverse and longitudinal dry-fibre stiffening elements, each of said transverse and longitudinal stiffening ele- ments comprising a plurality of layers;

c) arranging in a coordinated manner the support blocks alongside each other on a second unit, so as to form a dry-fibre stiffening grid with the transverse and longitudinal dry-fibre stiffening elements in contact with each other;

d) - turning over the second unit with the support blocks and arranging it on the first unit so as to position the dry-fibre stiffening grid on the dry-fibre laminate of the panel, so as to obtain a dry-fibre panel structure, or, alternatively,

- turning over the first unit and arranging it on the second unit with the support blocks so as to position the dry-fibre laminate of the panel on the dry-fibre stiffening grid and thus ob- tain a dry-fibre panel structure; and

e) subjecting the dry-fibre panel structure to vacuum resin infusion and subsequent heat treatment in an oven so as to obtain polymerization of the resin infused in the panel structure.

By means of the method according to the invention it is possible to produce, using dedicated equipment, panels such as wing and fuselage panels by means of a liquid infusion process which incorporates in the panel a stiffening system of the "grill" type namely comprising integrated intersecting stringers and/or spars with shear ties/ribs. It is thus possible to produce highly integrated structures, resulting in an overall reduction in weight and cost due to the elimination of the connection members and the thicker dimensions needed to strengthen the structure in the perforated zones, said connection members and holes being provided in conventional structures for connecting together the panel components.

Using the technology according to the present invention the manufacture of a fuselage panel (representing a section of the cockpit and therefore a surface with high curvature) has been completed and a wing panel (representing a bottom panel of a wing box of a UAV aircraft and therefore of notable length) is currently being manufactured.

Based on the production of a certain number of small-scale flat parts (using a scale-up approach for the technology) and the two aforementioned demonstration models it has been pos- sible to demonstrate the impact which the manufacturing process according to the present invention may have on the design factor of aeronautical components and the opportunities which it may offer. The reduction in costs may in fact be favoured by a more fully integrated struc- tural design with a consequent reduction in the number of parts.

Particular embodiments form the subject of the dependent claims, the contents of which are to be understood as forming an integral part of the present description.

Further characteristic features and advantages of the invention will be described more fully in the detailed description below of a number of embodiments thereof, provided by way of a non- limiting example, with reference to the attached drawings in which: Figure 1 is a perspective view which shows a small-scale flat panel manufactured using a method according to the invention;

Figure 2 is a perspective view which shows a step during manufacture of the panel structure according to Figure 1 , in which support blocks are assembled on a coordination unit;

Figure 3 is a plan view which shows a step during manufacture of the panel structure according to Figure 1 in which a vacuum bag is prepared;

Figure 4 is a perspective view which shows a wing panel manufactured using a method according to the invention;

Figure 5 is a perspective view showing support blocks used for manufacture of the panel structure according to Figure 4;

Figure 6 is a perspective view showing the support blocks of Figure 5 partially assembled and provided with transverse and longitudinal stiffening elements for the panel structure according to Figure 4;

Figure 7 is a perspective view which shows the support blocks of Figures 5 and 6 assembled so that the respective transverse and longitudinal elements are connected together to form a stiffening grid for the panel structure according to Figure 4;

Figure 8 is a perspective view which shows a flat wing panel arranged on a coordination unit provided with support blocks which support transverse and longitudinal stiffening elements for the panel structure;

Figure 9 is a perspective view which shows the coordination unit according to Figure 8 joined together with a unit for moving and rotating the coordination unit;

Figure 10 is a perspective view showing the coordination unit according to Figure 8 turned over and positioned on a curing unit containing a skin for the panel structure; Figure 11 is a perspective view which shows the curing unit containing the support blocks positioned on the skin for the panel structure;

Figure 12 is a perspective view which shows a curved panel manufactured using a method according to the invention;

Figure 13 is a perspective view showing a curing unit for manufacturing the panel structure according to Figure 12;

Figure 14 is a perspective view which shows a coordination unit and support blocks for manufacturing the panel structure according to Figure 12; and

Figures 15-17 are perspective views of a further panel which may be manufactured us- ing a method according to the invention.

With reference to Figure 1, this shows a small-scale panel of composite material denoted overall by 10. The panel 10 shown in Figure 1 comprises an outer lining or skin 11 formed by a plurality of layers of composite material consisting of fibre enclosed in a resin matrix. The panel 10 also comprises a pair of longitudinal stiffening elements 13 (known as stringers) which are fixed to the skin 11. Each of the longitudinal stiffening elements 13 is formed by a plurality of layers of composite material consisting of fibre enclosed in a resin matrix. In the example shown each longitudinal stiffening element 13 has a T-shaped cross-section and is formed by a pair of parts with an L-shaped cross-section joined together by means of respective flanges arranged in contact with each other.

The panel 10 also comprises a pair of transverse stiffening elements (shear ties) 15 which are fixed to the skin 11 and which interconnect the longitudinal stiffening elements 13. Each of the transverse stiffening elements 15 is formed by a plurality of layers of composite material consisting of fibre enclosed in a resin matrix. In the example shown each transverse stiffening element 15 has a T-shaped cross-section and is formed by a pair of parts with an L-shaped cross-section joined together by means of respective flanges arranged in contact with each other. For the purposes of the present invention, both the element with a T-shaped cross-section and the two L-shaped parts which form it may be understood as being stiffening elements (transverse or longitudinal, as appropriate). In any case the invention is not limited to the particular cross-sections described here.

For the purposes of the present invention the terms "longitudinal" and "transverse" are defined with respect to a direction in the surface of the skin of the panel structure, regarded randomly as being the longitudinal direction. The connection between longitudinal stiffening elements 13 and transverse stiffening elements 15 creates a stiffening grid for the skin 11. This connection is ensured by end parts of the transverse elements 15 which are configured to adhere to complementary surfaces (on the "head" or the "shank" of the T-shaped cross-section) of the longitudinal elements 13. A method for manufacturing the aforementioned panel 10 involves the steps described herein- below.

On a first unit, which will also be referred to below as curing unit (see in this connection the following embodiments) a dry-fibre laminate intended to form the skin 11 of the panel 10 is formed. The dry-fibre skin is made in such a way as to comprise a plurality of dry-fibre layers. This step envisages then the preparation and stratification planar-wise of the reinforcing layers (i.e. superimposition of the dry layers by positioning, depending on the design of the component, the layers with the fibres oriented as per the design and with the addition, where necessary, of additional reinforcing layers) using commercially available materials such as Non- Crimp Fabrics (NCF), dry tow placement materials or flat braiding manufactured by operators in the sector with predetermined geometric characteristics (depending on the requisites of the aircraft to be constructed) and physical characteristics (type of reinforcement fibre, areal weight of the fibres in the different orientations required, etc.). Furthermore transverse dry-fibre elements - with an L-shaped cross-section - are made, these elements being intended to form the transverse stiffening elements 15 with a T-shaped cross- section. The transverse dry-fibre stiffening elements are made so as to comprise a plurality of dry-fibre layers. This step envisages then the preparation and stratification of the reinforcing layers (i.e. superimposition of the dry layers by positioning, depending on the design of the component, the layers with the fibres oriented as per the design and with the addition, where necessary, of additional reinforcing layers) using commercially available materials such as Non-Crimp Fabrics (NCF), dry tow placement materials or flat braiding manufactured by operators in the sector with predetermined geometric characteristics (depending on the requisites of the aircraft to be constructed) and physical characteristics (type of reinforcement fibre, areal weight of the fibres in the different orientations required, etc.). The transverse dry-fibre stiffening elements may be stratified on a flat unit and then subjected to hot-forming on support blocks, which will be described below. In alternative embodiments, the transverse dry-fibre stiffening elements may be formed directly on the support blocks.

Furthermore longitudinal dry-fibre elements - with an L-shaped cross-section are made, these elements being intended to form the longitudinal stiffening elements 13 with a T-shaped cross- section. The longitudinal dry-fibre stiffening elements are made so as to comprise a plurality of dry-fibre layers. This step involves therefore the preparation and stratification of the reinforcing layers (i.e. superimposition of the dry layers by positioning, depending on the design of the component, the layers with the fibres oriented as per the design and with the addition, where necessary, of additional reinforcing layers) using commercially available materials such as Non-Crimp Fabrics (NCF), dry tow placement materials or flat braiding manufactured by operators in the sector with predetermined geometric characteristics (depending on the requisites of the aircraft to be constructed) and physical characteristics (type of reinforcement fibre, areal weight of the fibres in the different orientations required, etc.).

The longitudinal dry-fibre stiffening elements may be stratified on a flat unit and then subjected to hot-forming on support blocks, which will be described below. In alternative embodiments, the longitudinal dry-fibre stiffening elements may be formed directly on the support blocks.

Figure 2 shows the aforementioned support blocks - indicated by the reference numbers 20 and 30 - arranged in a coordinated manner on a unit 40, referred to below as coordination unit. These support blocks 20 and 30 have a polygonal form in a plan view and therefore comprise a plurality of perimetral sides. The geometrical form of the support blocks 20, indicated below as central blocks, is defined by the volume situated between two longitudinal stiffening elements and two transverse stiffening elements. Two lateral blocks 30 must be added to the aforementioned set of central blocks along the two longitudinal outer stiffening elements which extend along the entire length of the panel, there being no interruptions due to the presence of the transverse stiffening elements. In the example shown in Figure 2 three central support blocks 20 with a square or rectangular form and two lateral support blocks 30 with a rectangular form are provided. The blocks 20 and 30 represent the most important elements in the chain of units for formation of the integrated intersections of the panel. They determine the final form of the part, ensuring that strict geometrical tolerances are satisfied. In view of these requirements, as well as their large num- ber and the need to move the blocks 20 and 30 during manufacture, these blocks are conveniently made of light material with a low thermal expansion coefficient, for example composite material. The low TEC (thermal expansion coefficient) is such as to avoid undesirable movements of the stiffening elements 13 and 15 during heating in the curing cycle. The low weight of each block 20, 30 allows easy and ergonomic handling, for example using high-capacity suction grips in the case where the blocks are manually handled.

In order to ensure coordinated positioning and where necessary securing of the support blocks 20 and 30 on the coordination unit 40, said unit may be provided with connection means and/or securing means and/or indexing means which are known per se (not shown).

Before positioning the blocks 20 and 30 on the coordination unit 40 and/or during positioning of these blocks on the unit 40, positioning/forming of the stiffening elements 13 and 15 on the support bocks 20 and 30 is performed. In particular, rxjsitioning/forming is performed so that the transverse and longitudinal dry-fibre stiffening elements - with an L-shaped cross-section - are arranged at the top of the support blocks 20, 30, along the perimetral sides thereof. According to a possible procedure, the central support blocks 20 are arranged alongside each other so that respective perimetral sides are arranged adjacent and respective transverse dry- fibre reinforcing elements are in contact with each other along the direction of their length, so as to form the transverse elements with a T-shaped cross-section. In this way the central support blocks 20 arranged alongside each other form a strip of support blocks. This strip may be formed directly on the coordination unit 40 or formed separately and then transferred onto the coordination unit 40. Respective longitudinal dry-fibre reinforcing elements, with an L-shaped cross-section, are then arranged on opposite longitudinal sides of the strip of support blocks, so that these longitudinal dry-fibre reinforcing elements are in contact with the respect ends of all the transverse dry-fibre reinforcing elements of the strip of support blocks. The positioning forming of the longitudinal L-shaped elements on the strip of support blocks may be performed before or after positioning of the strip of support blocks 20 on the coordination unit 40. Finally, by means of the coordination unit 40, the strip of support blocks is arranged alongside the lateral support blocks 30, each having a respective longitudinal dry-fibre reinforcing element with L-shaped cross-section on a respective longitudinal side, so that each of the longitudinal sides of the strip of support blocks 20 is adjacent to the longitudinal side of the respective lateral support block 30 and respective longitudinal dry-fibre reinforcing elements - with an L- shaped cross-section - are in contact with each other along the direction of their length, so as to form the longitudinal elements with a T-shaped cross-section.

According to another possible procedure, one of the lateral support blocks 30 may be first positioned on the coordination unit 40, followed by arrangement of the strip of central support blocks 20 and finally the other lateral support block 30.

By positioning the support blocks 20, 30 on the coordination unit 40 the result is therefore achieved that the transverse and longitudinal dry-fibre stiffening elements mounted on these support blocks form a stiffening grid with the transverse and longitudinal dry-fibre stiffening elements in contact with each other.

After formation of the dry-fibre stiffening grid of longitudinal and transverse elements, positioning/indexing of the coordination unit 40 on the dry-fibre skin laminated on the curing unit is performed. For this purpose, the coordination unit 40 with the support blocks 20, 30 is turned over and arranged on the curing unit so as to position the dry-fibre stiffening grid on the dry-fibre skin of the panel and thus obtain a dry-fibre panel structure. In this structure, parts of the stiffening grid (in the example shown, the "heads" of the T-shaped cross-sections of the stiffening elements 13 and 15) come into contact with the surface of the skin 11. The coordination unit 40 is then removed, leaving the entire dry-fibre panel structure on the curing unit.

According to an alternative embodiment, a unit supporting the skin may be turned over and arranged on the coordination unit 40 with the support blocks 20 and 30 so as to position the dry- fibre skin of the panel on the dry-fibre stiffening grid and thus obtain a dry-fibre panel structure. The upper unit is then removed, leaving the entire dry-fibre panel structure on the coordination unit (used in this case also as a curing unit).

Thereafter (irrespective as to whether the stiffening grid has been positioned on the skin or vice versa) the dry-fibre panel structure is subjected to vacuum resin infusion and subsequent heat treatment in an oven so as to obtain polymerization of the resin infused in the panel structure.

For this purpose a vacuum bag is made (not shown) using materials and components known per se in the sector such as, starting from the dry-fibre part, separator film, breather fabric, bag film sealed on the edges of the polymerization unit and fixing of vacuum valves on the same bag film. Moreover channels are positioned for supplying and discharging the liquid resin during infusion in the oven. For example, Figure 3 shows a diagram of the structure produced in Figure 2, where the resin supply channels IC and resin discharge channels OC can also be seen. The discharge channels OC are positioned in areas where the two L-shaped parts of each stiffening element are joined together. The discharge channels OC have ends which emerge in respective areas of intersection between transverse stiffening elements 15 and longitudinal stiffening elements 13.

Once the vacuum bag has been formed, the component may thus be subjected to a dedicated infusion and polymerization cycle (at temperature without pressure).

Figure 4 shows a panel, for example a wing panel, which is more complex than that shown in Figure 1 , but may be likewise made using a method similar to that described above. This panel, indicated by 10', comprises a greater number of transverse stiffening elements 15 and longitudinal stiffening elements 13 than those of the panel 10 shown in Figure 1. While each of the longitudinal elements 13 is formed by a single part, without interruptions along its length, each of the transverse elements 15 is formed by a series of segments which interconnect separate longitudinal elements 13.

As shown in Figures 5 and 6, there is consequently a greater number of central support blocks 20 which are combined to form several strips of support blocks 21. Each strip of support blocks 21 is arranged alongside at least one other strip of support blocks 21 so as to have respective longitudinal sides arranged adjacent and respective longitudinal dry-fibre reinforcing elements - with an L-shaped cross-section - in contact with each other along the direction of their length, so as to form the longitudinal elements with T-shaped cross-section. Figure 7 shows the composition of the support blocks 20, 30 for forming the dry-fibre stiffening structure of the panel 10' according to Figure 4, comprising longitudinal stiffening elements (stringers) and transverse stiffening elements 15 (ribs). As regards the rest, the method for manufacturing the panel 10' in Figure 4 is substantially identical to that of the panel 10 shown in Figure 1 so that reference may be made to the description of said panel provided above.

Figure 8 shows a coordination unit CT comprising a plate 40' on which the coordinated positioning of the support blocks 20, 30 may be performed in order to create the stiffening grid of the panel 10' shown in Figure 4. For this purpose the unit CT may be provided with connection means and/or securing means and/or indexing means known per se.

Figure 9 shows an apparatus A configured to receive the coordination unit CT and perform overturning and positioning thereof on top of a curing unit PT (Figures 10 and 11), on which the dry-fibre skin 11 has been initially laminated. For correct positioning of the coordination unit CT on the curing unit PT, the coordination unit is provided with centring elements CM which engage with corresponding parts CC formed on the curing unit PT. Once the coordination unit CT has been removed, the vacuum bag may be formed on the curing unit PT so as to subject the dry-fibre component to the following infusion and polymerization treatment in an oven.

Figure 12 shows a panel, for example a fuselage panel, formed by a pair of curved panels which may be made using a method similar to that described above. The curved panels, indi- cated by 10", differ from those described above in that they have a curvature in the transverse direction. Each panel 10" comprises therefore a skin 11 and rectilinear longitudinal stiffening elements 13 (stringers) and curved transverse stiffening elements 15 (shear ties).

Figure 13 shows a curing unit PT" comprising a curved surface LS provided for lamination of the dry-fibre skin. Figure 14 shows a coordination unit CT" as well as central support blocks 20" and lateral support blocks 30" for constructing the stiffening grid formed by the stiffening elements 13 and 15. As can be seen in Figure 14, the central support blocks 20 have a trapezoidal form in plan view and a curved upper surface such as to match substantially the profile of the curved surface on which the skin 11 is positioned. As regards the rest, the method for manufacturing the panel 10" in Figure 12 is substantially identical to that of the panel 10 shown in Figure 1 so that reference may be made to the description of said panel provided above.

Figures 15, 16 and 17 show how the design is applicable also to those cases not covered by Figures 4 and 12 where the configuration of the wing and fuselage panels is respectively such that the longitudinal and transverse stiffening elements (13 and 15) are interrupted. In particular in the case of the longitudinal elements 13 it is possible to consider the possibility of designing the latter with an overall length which is smaller than the longitudinal dimension of the panel, while in the case of the transverse elements 15 it is possible for the single sections form- ing them to be interrupted with respect to the adjacent longitudinal element 13. This is made possible by the geometrical configuration of the support blocks which may comprise plugs and corresponding suitably machined recesses.

Different experiments for application of the present invention were carried out. The results ob- tained hitherto are particular encouraging because:

the quality of the manufactured components is optimum both in terms of absence of internal porosity and in particular as regards the strict dimensional tolerances which they are able to satisfy (the absence of pressure during curing in fact does not generate excessive variations in thickness and/or deformations of the composite parts);

it is possible to obtain the profiles of the longitudinal and transverse stiffening elements which are interrupted with a theoretical clean edge, already during manufacture owing to the prior cutting of the dry pre-forms, with a consequent advantage in terms of component manufacturing time and costs since the additional step of material removal performed in numerical- control machining centres is avoided;

the manufacturing cost is much lower compared to conventional pre-preg processes because the raw material is more convenient, it does not require storage of the dry fibre in a refrigerator, larnination is more rapid, curing is performed without pressure and the high integration reduces the number of parts;

the cost of assembly is much lower compared to conventional processes owing to the absence of boring operations and installation of connecting members to be performed on the skins. For this particular reason the configuration/process proposed is particularly suitable for applications involving wing components with fuel-wetted surfaces.